Conventional methods of forming stiffened composite structures may involve assembling uncured composite parts to form an uncured detail assembly of a stiffener. The composite parts may include a flange and a web interconnected by a bend radius. The detail assembly may also include a radius filler which may be encapsulated by a base laminate. The base laminate may constrain the flanges of the composite parts against spring-in which may occur due to resin shrinkage (e.g., chemical shrinkage) during curing of the detail assembly. In addition, the base laminate may also constrain the flanges against thermally-induced spring-in which may occur as the detail assembly cools down from the cure temperature to room temperature. The cured detail assembly may be co-bonded to an uncured layup such as a skin panel as a means to stiffen the skin panel.
Unfortunately, the constrainment of the flanges against spring-in may result in relatively high through-thickness residual tension in the bend radii and in the radius filler. The through-thickness residual tension may be locked into the detail assembly during co-bonding, and may result in undesirable effects on the stiffened skin panel such as a reduction in load-carrying capability. The magnitude of residual stress due to resin shrinkage and thermally-induced spring-in may generally increase with increasing laminate thickness. As a result, laminate thicknesses must be further increased so that the skin panel is capable of carrying the design loads. For weight-sensitive structures such as aircraft, an increase in laminate thickness may correspond to an increase in the structural mass of the aircraft which may have a detrimental effect on aircraft performance such as climb rate, payload-carrying capability, range, and/or fuel efficiency.
As can be seen, there exists a need in the art for a method of reducing residual stress in a composite assembly.